Electric field sensor with sensitivity-attenuating ground ring

ABSTRACT

An electric field sensor includes an insulating substrate, a plurality of electrodes, an insulator, a plurality vias, and a ground ring. The electrodes are disposed on the substrate. The insulator is disposed over the electrodes. The vias are coupled to the electrodes and extend through the substrate at a right angle to the electrodes. The ground ring is disposed around the electrodes and the vias and is configured to attenuate a sensitivity of the sensor to electric fields outwards of the ground ring.

RELATED APPLICATIONS

This patent application is a continuation of U.S. patent applicationSer. No. 15/571,588, filed Nov. 3, 2017, now U.S. Pat. No. 10,940,952,entitled “Apparatus and Method of Monitoring for In-Flight AircraftEngine Ice Crystal Accretion”, which is a § 371 of International PatentApplication Serial No. PCT/CA2016/050517, filed May 5, 2016, whichclaims the benefit of the filing date of U.S. Patent Application Ser.No. 62/156,891, filed May 5, 2015, the contents of all of which areincorporated herein by reference.

FIELD

This patent application relates to an electric field sensor. Inparticular, this patent application relates to an electric field sensorthat may be used to monitor for the in-flight accretion of ice crystalson aircraft engine surfaces.

BACKGROUND

Aircraft may be exposed to various atmospheric icing conditionsin-flight, such as small water droplets less than 25 um diameter (e.g.clouds, fog) and super-cooled large water droplets greater than 25 umdiameter (e.g. freezing rain, freezing drizzle). Ice accretion resultingfrom exposure to these atmospheric icing conditions, in and around thewarm surfaces of an aircraft engine, such as the cowling and/or airintake ducts of turbofan engines, the carburetor mouth and/or throttlebody of piston engines, and even the internal surfaces of the engine,can cause engine damage and/or a sudden loss in engine output power andaircraft stability. Therefore, many modern aircraft engines incorporatecountermeasures, such as heating systems, to reduce the likelihood orextent of engine ice accretion on the external and internal enginesurfaces. However, over the past 20 years there have been over 100reported cases of aircraft engine power loss as a result of iceaccretion on engine surfaces.

SUMMARY

This patent application describes an electric field sensor that includesan insulating substrate, a plurality of electrodes, an insulator, aplurality vias, and a ground ring.

The electrodes are disposed on the substrate. The insulator is disposedover the electrodes.

The vias are coupled to the electrodes and extend through the substrateat a right angle to the electrodes.

The ground ring is disposed around the electrodes and the vias and isconfigured to attenuate a sensitivity of the sensor to electric fieldsoutwards of the ground ring.

The plurality of electrodes may include a plurality of first electrodesand a plurality of second electrodes that are interleaved andnon-contacting with the plurality of first electrodes. The firstelectrodes are not in contact with each other, the second electrodes arenot in contact with each other, and the plurality of first electrodesare disposed parallel to the plurality of second electrodes on thesubstrate.

The plurality of vias may include a first via portion and a second viaportion. The vias of the first via portion are coupled to the pluralityof first electrodes, and the vias of the second via portion are coupledto the plurality of second electrodes.

The insulator may be configured to be mounted substantially flush withthe surface of an aircraft engine.

The insulator may be aerodynamically-matched to the surface of theaircraft engine.

The insulator may be thermally-matched to the surface of the aircraftengine.

Proposed aircraft engine surfaces include the cowling of a turbofanengine, the air intake duct of a turbofan engine, the carburetor mouthof a piston engine, the throttle body of a piston engine, and aninternal engine surface.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing aspects will now be described, by way of example only,with reference to the accompanying drawings, in which:

FIG. 1 is a schematic view of a matter accumulation monitoring system,depicting the electric field sensor unit and a sensor processing unit;

FIG. 2 a is a perspective view of the electric field sensor;

FIG. 2 b is a top plan view of the electric field sensor;

FIGS. 2 c and 2 d are transverse cross-sectional views of the electricfield sensor;

FIG. 3 is a schematic view of the structure of the sensor processingunit; and

FIG. 4 is a flow-chart depicting a method of operation of the matteraccumulation monitoring system.

DETAILED DESCRIPTION

1. Matter Accumulation Monitoring System: Overview

Turning now to FIG. 1 , there is shown a matter accumulation monitoringsystem, denoted generally as 100, comprising one or more electric fieldsensor units 200, a sensor processing unit (SPU) 300, and a wiringharness 102 interconnecting the electric field sensor units 200 and thesensor processing unit (SPU) 300.

As shown, the electric field sensor units 200 are preferably separateand distinct from the SPU 300. Although the functionality of the SPU 300may also be incorporated into each of the electric field sensor units200, it should be understood that maintaining the electric field sensorunits 200 separate and distinct from the SPU 300 reduces the physicalsize of the electric field sensor units 200 and allows the electricfield sensor units 200 to be better aerodynamically-matched andthermally-matched to the associated aircraft surfaces/engine surfaces.

2. Electric Field Sensor Unit

Each electric field sensor unit 200 is disposed proximate a surface ofan aircraft, at a respective region thereof, to monitor the accumulationof matter on the respective region of the aircraft surface while theaircraft is in-flight.

The matter to be monitored may be ice, and one or more of the electricfield sensor units 200 may be mounted on and thermally-coupled to anexterior surface of an aircraft at regions thereof that are prone to theaccumulation of ice while the aircraft is in-flight. Preferably, theelectric field sensor units 200 are aerodynamically-matched to theaircraft surfaces being monitored so as to not disturb the normalairflow and hence the normal buildup of ice on the associated aircraftsurfaces. Further, preferably the electric field sensor units 200 arealso thermally-matched to the associated aircraft surfaces so as toagain not interfere with the normal buildup of ice on the aircraftsurfaces.

More preferably, one or more of the electric field sensor units 200 ismounted to and thermally-coupled to an engine surface of an aircraft atregions of the engine that are prone to the accretion of ice crystalswhile the aircraft is in-flight. Typically an aircraft is fitted with aplurality of the electric field sensor units 200. It is expected that,through experimentation, the person of ordinary skill could determinethe appropriate placement of the electric field sensor units 200 todetect ice crystal accretion on engine surfaces while the aircraft is inflight.

Preferably, the electric field sensor units 200 areaerodynamically-matched to the engine surfaces being monitored so as tonot disturb the normal airflow and hence the normal buildup of icecrystals on the engine surfaces. Accordingly, the electric field sensorunits 200 may be flush-mounted with the respective engine surfaces. Asnon-limiting examples, one or more of the electric field sensor units200 may be flush-mounted with a cowling of a turbofan engine, an airintake duct of a turbofan engine, a carburetor mouth of a piston engine,a throttle body of a piston engine, or an internal engine surface.

Although conventional ice mitigation countermeasures have beensuccessful in limiting ice accretion on engine surfaces from exposure tosmall water droplets less than 25 um diameter (e.g. clouds, fog) andsuper-cooled large water droplets greater than 25 um diameter (e.g.freezing rain, freezing drizzle), the inventors have learned that fullyglaciated water (e.g. ice crystals) can accumulate on the heatedsurfaces of aircraft engines. In contrast to an aircraft wing where thewing surface remains cold while the aircraft is in flight, the inventorsbelieve that heat transfer from an engine surface while the aircraft isin flight initially causes the ice crystals to melt, but that rapid airflow over the engine surface subsequently cools the melt and causes theice crystals to adhere to the engine surface. Therefore, so as to againnot interfere with the normal buildup of ice on the engine surfaces,preferably the electric field sensor units 200 are alsothermally-matched to the associated engine surfaces in the sense thatthe field sensor units 200 absorb heat from (and introduce heat into)the airstream passing over the electric field sensor units 200 at thesame rate as the associated aircraft surface/engine surfaces.

As shown in FIG. 1 , preferably each electric field sensor unit 200includes a signal amplifier 204, and an electric field sensor 202coupled to the signal amplifier 204. Optionally, the electric fieldsensor unit 200 may also include a digital temperature sensor 203, and aheater (not shown).

Each electric field sensor 202 generates a time-varying local electricfield at the engine surface, and a current is produced in the electricfield sensor 202 resulting from the time-varying local electric field.The magnitude and phase of the resulting current varies with thecharacteristics of the material that is within the electric fieldestablished by the electric field sensor 202. Accordingly, the electricfield sensor(s) 202 together provide periodic data samples indicative ofthe accretion of ice crystals on the associated aircraft surface/enginesurface.

The temperature sensors 203 monitor the temperature of the associatedaircraft surfaces/engine surfaces and are typically used to calibratethe electric field sensor 202. Alternately, or additionally, thetemperature sensor(s) 203 may be used in the detection of ice crystalson the associated aircraft surfaces/engine surfaces.

As shown in FIGS. 2 a, 2 b and 2 c , each electric field sensor 202comprises an insulating substrate 206, a plurality of electrodes 208disposed on the substrate 206, and a plurality of vias 210 coupled tothe plurality of electrodes 208 and extending downwardly through thesubstrate 206. Preferably, the substrate 206 comprises a non-conductivematerial, such as ceramic, although other non-conductive materials maybe used. The electrodes 208 are typically formed on the substrate 206using conventional printed circuit board or integrated circuitmanufacturing techniques. The electrodes 208 extend across the topsurface of the substrate 206 in a substantially parallel fashion, suchthat the electrodes 208 do not contact one another on the top surface ofthe substrate 206.

The electrodes 208 are segregated into a first electrode portion 208 a,and a second electrode portion 208 b. The electrodes 208 of the firstelectrode portion 208 a are not in contact with one another and extendfrom one end 212 a of the substrate 206, and the electrodes 208 of thesecond electrode portion 208 b are not in contact with one another andextend from the opposite end 212 b of the substrate 206. The electrodes208 of the first electrode portion 208 a are interleaved with and aredisposed substantially parallel to the electrodes 208 of the secondelectrode portion 208 b in the centre region 214 of the top surface ofthe substrate 206.

Typically, each via 210 comprises a plated through-hole extending fromone end of a respective electrode 208, through the substrate 206,towards the bottom surface 216 of the electric field sensor unit 200.Alternately, the vias 210 may be provided as conductive traces or wiresextending in a similar manner. The vias 210 are segregated into a firstvia portion 210 a and a second via portion 210 b. The vias 210 of thefirst via portion 210 a are coupled to the first electrode portion 208a, and the vias 210 of the second via portion 210 b are coupled to thesecond electrode portion 208 b.

Each via 210 is connected to a respective electrode 208 adjacent therespective end 212, and extends at a right angle from the electrode 208through the substrate 206, from the top surface of the electric fieldsensor unit 200 towards the bottom surface 216 of the electric fieldsensor unit 200. Preferably, the electric field sensor 202 also includesa ground ring 218, disposed on the substrate 206 laterally outwards fromthe electrodes 208 and the vias 210. The ground ring 218 may beconnected to electrical grounds of the signal amplifier 204. With thisconfiguration, the sensitivity of the electric field sensor 202 toelectric fields outside the centre region 214 is less than conventionalelectric field sensors.

Each via 210 of the first via portion 210 a terminates at a firstconductive plate 220 that is embedded within the substrate 206 of theelectric field sensor unit 200. Similarly, each via 210 of the secondvia portion 210 b terminates at a second conductive plate (not shown)that is embedded within the substrate 206 (but separate from the firstconductive plate 220).

The electric field sensor 202 preferably also includes a topelectrically-insulating layer 222 disposed over the electrodes 208, anda bottom electrically-insulating layer 224 disposed over the bottomsurface 216 of the electric field sensor unit 200. As discussed,preferably the electric field sensor units 200 areaerodynamically-matched to the aircraft surfaces/engine surfaces beingmonitored. Accordingly, preferably the top insulating layer 222 isaerodynamically-matched to the shape and contour of the associatedaircraft surface/engine surface so as to not disturb the normal airflowand hence the normal buildup of ice crystals on the aircraftsurface/engine surface. Typically, the top insulating layer 222 isdisposed substantially flush with the associated aircraft surface/enginesurface. As discussed, suitable engine surfaces include, but are notlimited to, the cowling of a turbofan engine, the air intake duct of aturbofan engine, the carburetor mouth of a piston engine, the throttlebody of a piston engine, and an internal engine surface.

As discussed, preferably the electric field sensor units 200 arethermally-coupled and thermally-matched to the aircraft surfaces/enginesurfaces being monitored. Accordingly, so as to further not interferewith the normal buildup of ice on the aircraft surface/engine surface,preferably the top insulating layer 222 is thermally-matched to theassociated aircraft surface/engine surface such that the top insulatinglayer 222 absorbs heat from (and introduces heat into) the airstreampassing over the insulating layer at the same rate as the associatedaircraft surface/engine surface. Typically, the insulating layers 222,224 each comprises a ceramic or glass, although the ceramic alumina ispreferred due to its hardness and thermal conductivity

Preferably, the signal amplifier 204 is incorporated into the electricfield sensor unit 200 below the electric field sensor 202 and isconnected to the vias 210 of the electric field sensor 202 via theconductive plates 220. First and second electrically-conductive pins 226a, 226 b extend from the signal amplifier 204 and out the bottom surface216 of the electric field sensor unit 200.

As will be explained, the sensor processing unit (SPU) 300 appliestime-varying analog voltage signals to the first and second conductivepins 226 a, 226 b via the wiring harness 102. The signal amplifier 204receives the time-varying analog voltage signals via the conductive pins226 a, 226 b, amplifies the analog voltage signals, and applies theamplified analog voltage signals to the electrodes 208 of the electricfield sensor 202 via the vias 210. Preferably, the analog voltage signalapplied to the first electrode portion 208 a is complementary (i.e. 180degrees out of phase) with the analog voltage signal applied to thesecond electrode portion 208 b.

The amplified analog voltage signals cause the electrodes 208 to apply atime-varying electric field to the aircraft surface/engine surface thatis proximate the electric field sensor 202. The time-varying electricfield causes an analog current to be produced in the electrodes 208 ofthe electric field sensor 202. The signal amplifier 204 detects theresulting analog current via the vias 210 of the electric field sensor202, generates an analog voltage signal from the resulting current, andapplies the generated analog voltage signal to the first and secondconductive pins 226 a, 226 b. The SPU 300 detects the generated analogvoltage signal at the first and second conductive pins 226 a, 226 b viathe wiring harness 102.

Preferably, the temperature sensor 203 is incorporated into theassociated electric field sensor 202 proximate the top surface of thesubstrate 206, and the electric field sensor 202 includes a via 210 thatextends through the substrate 206 from proximate the top surface towardsthe bottom surface 216 of the electric field sensor unit 200. A thirdelectrically-conductive pin 222 c extends from the via 210 of thetemperature sensor 203 and out the bottom surface 226 of the electricfield sensor unit 200. Alternately, each temperature sensor 203 may bedisposed in a substrate that is separate from the substrate 206 of theelectric field sensor unit 200.

The temperature sensor 203 generates a serial digital output signal thatincludes temperature measurement samples of the ambient temperatureproximate the electric field sensor 202, and applies the generateddigital temperature measurement signal to the third conductive pin 226c. The SPU 300 receives the temperature measurement samples, that areoutput at the third conductive pin 226 c, via the wiring harness 102.

Preferably, the heater is also incorporated into the associated electricfield sensor 202 proximate the top surface of the substrate 206, and isconnected to the SPU 300 via the wiring harness 102. As will beexplained, the SPU 300 uses the heater to melt ice that may haveaccumulated proximate an electric field sensor unit 200. However, so asto avoid frustrating the detection of ice at an electric field sensorunit 200 (by interfering with the thermal matching between the electricfield sensor unit 200 and the aircraft surfaces/engine surfaces beingmonitored), the heater is otherwise typically inactive while theaircraft is in flight.

3. Sensor Processing Unit

The sensor processing unit (SPU) 300 is typically disposed within theaircraft cockpit, and is coupled to the electric field sensor units 200via the wiring harness 102. As shown in FIG. 3 , preferably the SPU 300includes a sensor monitor 302, and an analog signal generator 320, andan A/D converter 322.

The sensor monitor 302 comprises a non-volatile non-transient memory304, a monitor interface 306, and a central processing unit (CPU) 308that is coupled to the non-volatile memory 304 and the monitor interface306. The monitor interface 306 may interface the SPU 300 with theinstrumentation of the aircraft cockpit to thereby provide pilots with asubstantially-real time assessment of the accumulation of matter on theaircraft surfaces/engine surfaces.

The non-transient memory 304 may be provided as an electronic memory, amagnetic disc and/or an optical disc, and may include a signaturesdatabase 350 of one or more predetermined matter accumulation profiles.Each predetermined matter accumulation profile is associated with aparticular characteristic (e.g. thickness, matter type) of the matteraccumulated, and comprises a corresponding time-series of complexadmittance values and optionally temperature values. Typically eachpredetermined matter accumulation profile includes a time-series ofcomplex admittance values and optionally temperature values for aparticular matter accumulation characteristic while the aircraft is inflight. The time-series of complex admittances and temperatures (ifincluded) in each matter accumulation profile may be predeterminedexperimentally and/or via computer modelling, and may be typicallystored in the memory 304 prior to installation of the SPU 300 in theaircraft.

The non-transient memory 304 also stores processing instructions for theSPU 300 which, when executed by the CPU 308, may define a signaturemonitor 310 and a signal processor 330. The signature monitor 310commands the analog signal generator 320 to generate and apply atime-varying analog voltage signal to the signal amplifier 204 of eachof the electric field sensor units 200. As discussed, the signalamplifier 204 amplifies the received analog voltage signal, and appliesthe amplified analog voltage signals to the electrodes 208 of theelectric field sensor 202. The amplified analog voltage signal cause theelectrodes 208 to apply a time-varying electric field to the aircraftsurface/engine surface that is proximate the electric field sensor 202.The time-varying electric field causes an analog current to be producedin the electrodes 208 of the electric field sensor 202. The signalamplifier 204 detects the resulting analog current, and generates ananalog voltage signal from the resulting current.

The A/D converter 322 periodically digitizes, over a measurement timespan, the analog voltage signals generated by the signal amplifier 204of the electric field sensor(s) 202 and provides the signal processor330, in substantially real-time, with the digitized version of theanalog voltage signals (hereinafter the “digitized current measurementsamples”).

The signal processor 330 uses the digitized current measurement samplesfrom the A/D converter 322, and optionally the digital temperaturemeasurement samples from the temperature sensors 203, to create atime-series of measurement data sets. Each measurement data set includesa magnitude measurement and a phase measurement. Preferably, the signalprocessor 330 derives the magnitude and phase measurements from thedigitized current measurement samples by referencing the magnitude andphase of the current produced in the electric field sensor(s) 202respectively to the magnitude and phase of the applied sensor voltagesignals. In effect, then, the magnitude and phase measurements arecomplex admittance measurements. However, for ease of reference, thecurrent magnitude and current phase measurements (referenced to theapplied sensor voltage) will be referred to hereinafter respectively asmagnitude measurements and phase measurements and collectively ascomplex admittance measurements.

As discussed, the temperature sensors 203 are typically used tocalibrate the electric field sensor 202. Therefore, for example, forcalibration, the signal processor 330 may use temperature samplesreceived from the temperature sensor(s) 203 to compute weight factors toapply to the complex admittance measurements.

In addition to the complex admittance measurements, each measurementdata set may optionally also include the digitized temperaturemeasurement sample that was taken when the associated digitized currentmeasurement sample was generated. Each measurement data set may alsoidentify the time at which the complex admittance measurements andtemperature measurements were taken.

The inventors have determined that the values for the complex admittancemeasurements, and also the variability (e.g. rate of change, range offluctuation) in those values between successive measurements, varieswith the characteristics of the matter accumulating on the aircraftsurface/engine surface. Therefore, to facilitate the detection of matteraccumulating on the aircraft surface/engine surface, and thedifferentiation between different matter accumulating on the aircraftsurface/engine surface, the signal processor 330 provides the signaturemonitor 310 with the time-series of the complex admittance measurements(and optionally temperature measurements). The signature monitor 310 isconfigured to use the time-series of complex admittance measurements(and optionally temperature measurements) to generate an assessment ofthe instantaneous accumulation of matter on the aircraft surface/enginesurface in substantially-real time.

The signature monitor 310 may generate the assessment by querying thesignatures database 350 with the received measurement data sets forcorresponding (identical or similar) predetermined matter accumulationprofiles. Alternately, the signature monitor 310 (or the signalprocessor 330) may be configured with a matter assessment algorithm thatgenerates the assessment of the instantaneous accumulation of matterfrom the measurement data sets.

The signature monitor 310 may then generate and transmit an alarm signalto the aircraft cockpit in real time if one or more of the receivedmeasurement data sets correlates with one of the predetermined matteraccumulation profiles, indicating that ice crystals have accreted on anaircraft surface/engine surface. Alternately, the signature monitor 310may generate and transmit an alarm signal to the aircraft cockpit inreal time if the matter assessment algorithm determines that icecrystals have accreted on an aircraft surface/engine surface.

As discussed above, the signature monitor 310 and the signal processor330 may each be implemented as a set of computer processinginstructions. However, the implementation of the signature monitor 310and the signal processor 330 is not so limited, but may each beimplemented instead in electronics hardware, such as a fieldprogrammable logic gate array (FPGA) or a complex programmable logicdevice (CPLD).

4. Matter Accumulation Monitoring System: Method of Operation

The method of operation of the matter accumulation monitoring system 100will now be described with reference to FIG. 4 .

At step S400, the matter accumulation monitoring system 100 appliestime-varying electric fields to the aircraft surfaces and/or enginesurfaces. To do so, the analog signal generator 320 of the SPU 300 mayapply a time-varying analog voltage signal to the electrodes 208 of theelectric field sensor(s) 202 (via the signal amplifier 204), with thevoltage signal applied to the first electrode portion 208 a being 180degrees out of phase with the voltage signal applied to the secondelectrode portion 208 b. The alternating electric fields are typicallyapplied to the aircraft surfaces/engine surfaces while the aircraft isin flight.

At step S402, the SPU 300 generates a time-series of measurement datasets, each measurement data set comprising a complex admittancemeasurement of the resulting current produced in the electric fieldsensor unit 200, and optionally the temperature of the aircraftsurfaces/engine surfaces. To do so, the A/D converter 322 mayperiodically sample the resulting current produced in the associatedelectric field sensor(s) 202, and the signal processor 330 may generatethe time-series of measurement data sets from the digitized currentmeasurement samples and optionally the temperature measurement samples.As noted above, the values in the measurement data sets, and thevariability in those values between successive measurement data sets,will vary in accordance with the characteristics of the matteraccumulating on the aircraft surfaces/engine surfaces.

At step S404, the SPU 300 generates an assessment of the instantaneousaccumulation of matter (e.g. ice crystals) on the aircraftsurfaces/engine surfaces. To do so, the signature monitor 310 may querythe signatures database 350 with the time-series of measurements fromone or more of the received measurement data sets, and generate theassessment from a correlation between the time-series of measurementsand the predetermined matter accumulation profiles stored in thesignatures database 350. As discussed above, each predetermined matteraccumulation profile is associated with a particular characteristic(e.g. thickness, matter type) of accumulated matter (e.g. ice crystals).

Where the time-series of received measurements correlates well with aparticular matter accumulation profile (the received measurement datasets are identical or similar to one or more of the predetermined matteraccumulation profiles), the SPU 300 may generate the assessment from thecharacteristics of the located matter accumulation profile.

However, typically an aircraft will be fitted with a plurality ofelectric field sensor units 200, each located at a respective aircraftsurface/engine surface. Accordingly, the SPU 300 may generate theassessment by querying the signatures database 350 with the measurementdata sets received from a plurality of the electric field sensor units200, and generate the assessment from a correlation between the receivedmeasurement data sets and a plurality of the predetermined matteraccumulation profiles. In this variation, the assessment may includecharacteristics from the various matter accumulation profiles.Alternately, the signature monitor 310 may generate the assessment byapplying the measurement data sets as inputs to a matter assessmentalgorithm.

The SPU 300 may thereafter transmit the results of the assessment to theaircraft cockpit for display on cockpit instrumentation. Alternately, oradditionally, the SPU 300 may initiate an automated control action (e.g.activate an alarm, invoke a reduction in the aircraft speed) inaccordance with a result of the correlation. For example, the SPU 300may activate an alarm if the signature monitor 310 determines that icecrystals in excess of a threshold amount have accreted on an aircraftsurface/engine surface).

After an excess ice accumulation has been detected, the SPU 300 mayactivate the heater of the associated electric field sensor unit(s) 200to thereby melt the ice accumulation at the respective electric fieldsensor unit(s) 200. During this phase, the SPU 300 may continue to usethe measurement data sets received from the electric field sensor units200 to generate an assessment of the instantaneous accumulation of icecrystals on the aircraft surfaces/engine surfaces, and may maintain theheater(s) active until the assessment indicates that the iceaccumulation proximate the respective electric field sensor unit(s) 200has been shed. Accordingly, the SPU 300 may periodically activate anddeactivate the heater(s) as the aircraft is in flight, in accordancewith the detected ice accumulation.

The SPU 300 may also correlate the measurement data sets received fromthe electric field sensor units 200 (while a heater is active and theice accumulation is being shed) with the matter accumulation profiles tofurther confirm the identity of the substance that had accumulatedproximate the electric field sensor unit(s) 200.

We claim:
 1. An electric field sensor comprising: an insulatingsubstrate; a plurality of electrodes disposed parallel to one another onthe substrate; an insulator disposed over the plurality of electrodes; aplurality vias coupled to the plurality of electrodes and extendingthrough the substrate at a right angle to the plurality of electrodes;and a ground ring disposed around the plurality of electrodes and theplurality of vias and configured to attenuate a sensitivity of thesensor to electric fields outwards of the ground ring; wherein theinsulator is configured to be mounted substantially flush with a surfaceof an aircraft engine.
 2. The sensor according to claim 1, wherein theplurality of electrodes comprises a plurality of first electrodes and aplurality of second electrodes interleaved and non-contacting with theplurality of first electrodes, wherein the first electrodes are not incontact with each other, and the second electrodes are not in contactwith each other, and the plurality of first electrodes are disposedparallel to the plurality of second electrodes on the substrate.
 3. Thesensor according to claim 2, wherein the plurality of vias comprises afirst via portion and a second via portion, the vias of the first viaportion are coupled to the plurality of first electrodes, and the viasof the second via portion are coupled to the plurality of secondelectrodes.
 4. The sensor according to claim 1, wherein the insulator isaerodynamically-matched to the surface of the aircraft engine.
 5. Thesensor according to claim 4, wherein the insulator is thermally-matchedto the surface of the aircraft engine.
 6. The sensor according to claim5, wherein the surface comprises one of a cowling of a turbofan engine,an air intake duct of a turbofan engine, a carburetor mouth of a pistonengine, a throttle body of a piston engine, and an internal enginesurface.